1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an industrial gas turbine engine turbine blade with platform cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The cooling of the blade platform in an industrial gas turbine engine is produced using convection cooling or film cooling. In the convection cooled platform, straight cooling holes formed within the platform with long length-to-diameter ratios are used. FIGS. 1 and 2 show this prior art blade platform cooling design using convection cooling holes. FIGS. 3 and 4 show the prior art blade platform cooling design using film cooling holes. The blade includes an airfoil section 11 extending from a platform 12 and a root section 13 with a cooling air supply channel 16. In FIG. 3, the platform is cooled using a number of film cooling holes 15 connected to a dead rim cavity 14 formed below the platform 12. In FIG. 2, the platform convection cooling holes are supplied from the cooling air supply channel 16.
The blade platform cooling designs of FIGS. 1 through 4 have several important design issues. Providing film cooling air for the entire blade platform requires a cooling air supply pressure from the dead rim cavity 14 to be higher than the peak blade platform external gas side pressure. This design induces a high leakage flow around the blade attachment region 13 and therefore causes a performance penalty. Using the long length-to-diameter ratio convection cooling holes that are drilled from the platform edge to the airfoil cooling supply channel 16 from the blade platform produces unacceptable stress levels at the airfoil cooling core and the platform cooling channels interface location, which therefore yields a low blade life. This problem is primary due to the large mass at the front and back ends of the blade root or attachment 13 which constrains the blade platform expansion. The cooling channels are also oriented transverse to the primary direction of the stress field which produces high stress concentrations in the cooling channels at the entrance location. Also, drilling the long cooling holes 23 along the platform axial length from the side will not cover the local hot spot 25 on the blade pressure side platform identified in FIG. 5 because of the angled cooling holes 22 on the suction side surface.
FIG. 5 shows a prior art turbine blade platform cooling circuit with long length-to-diameter cooling channels 21 on the suction side that feed smaller cooling channels 22 that branch off at an angle, and several long length-to-diameter cooling channels 23 that extend along the pressure side surface of the platform and that are parallel. These cooling channels are supplied from a dead rim cavity located below the platform and discharge onto the aft side edge of the platform.